This application is based on and claims the priority under 35 U.S.C. xc2xa7119 of German Patent Application 199 58 310.2, filed on Dec. 3, 1999, the entire disclosure of which is incorporated herein by reference.
The invention relates to a liquid flow rocket engine with a closed flow cycle. Such an engine is equipped with at least one main liquid flow supply line for each fuel component including fuel and oxygen components. The supply line or lines are connected to the combustion chamber. A turbo pump it provided in each liquid fuel component supply line.
Liquid fuel rocket engines of the type described above are normally equipped with a gas generator for driving the turbo pump or pumps in the fuel component supply lines. The gas generator is connected to at least one bypass fuel line to provide fuel for operating the gas generator. The exhaust gases from the exhaust gas generator are fed through a respective exhaust gas duct back into the main fuel supply line or lines.
Liquid fuel rocket engines are generally divided into two types, namely engines with a closed flow cycle and engines with an open flow cycle. Such engines are described in U.S. Pat. No. 5,404,715 relating to rocket engines. The engine with a closed flow cycle is referred to as xe2x80x9cdirect auxiliary flow cyclexe2x80x9d engine. Engines with an open flow cycle are referred to as xe2x80x9cparallel auxiliary flow cyclexe2x80x9d engines. Both types of engine operations according to the prior art have however certain disadvantages. Engines operating with an open flow cycle have a simpler construction particularly with regard to the interface between the main fuel supply line or lines and the combustion chamber. However, open cycle engines have a separate discharge for the exhaust gases of the gas generator whereby impulse losses are unavoidable. Such impulse losses adversely affect the thrust power of the rocket engine. Rocket engines with a closed flow cycle avoid such impulse losses, however closed flow cycle engines have the disadvantage that a complicated and expensive construction is required particularly of the interface between the main flow supply lines and the combustion chamber. The structural components in the interface must all be constructed with due regard to the configuration, size and position of all other components of the interface area which makes the construction and design involved and hence expensive.
In view of the above it is the aim of the invention to achieve the following objects singly or in combination:
to provide a liquid fuel rocket engine with a closed flow cycle which nevertheless has a simplified construction while maintaining the advantages of the closed flow cycle;
to improve the engine""s cavitation characteristics and to avoid evaporation of the liquid flow in the main liquid fuel supply line or lines;
to achieve a rapid mixing of the exhaust gases of the gas generator with the main fuel flow or at least one fuel component, such as liquid oxygen, while simultaneously avoiding an incomplete condensation of the gas generator exhaust gases; and
to facilitate the initial ignition when starting a closed flow cycle liquid fuel rocket engine.
According to the invention there is provided a liquid fuel rocket engine with a closed flow cycle, which engine is characterized by the combination of a combustion chamber with at least one main liquid fuel line for each liquid fuel component for supplying liquid fuel to the combustion chamber. A turbo pump is provided in each of the main liquid fuel supply lines. A gas generator for driving the turbo pump or pumps is connected to a first fuel supply byass line for feeding fuel or at least a fuel component, such as liquid oxygen, to operate the gas generator. The exhaust gas of the gas generator is fed through an exhaust gas duct to at least one of the liquid fuel supply lines to a point of the main liquid fuel supply line or lines upstrearm of a respective turbo pump or turbo pumps. In this context the term xe2x80x9cliquid fuel componentsxe2x80x9d encompasses the liquid fuel and the liquid oxygen.
According to the invention the exhaust gas duct feeds the gas generator exhaust gas into the main fuel supply stream in an area upstream of the turbo pumps, so that the respective exhaust gas duct or ducts are connected to the respective main fuel supply line at a location where their position is no longer critical to the construction of the interface between the fuel supply lines and the combustion chamber. This construction of the invention is contrary to the prior art in which the exhaust gas ducts are connected to the combustion chamber or to the injection head that leads into the combustion chamber for supplying exhaust gas into the main fuel flow. If only one turbo pump is provided in the main fuel supply line, the invention teaches that the exhaust gas duct or ducts are connected to that main fuel supply line upstream of the respective turbo pump. In case there are several turbo pumps arranged in series in the main fuel supply flow the invention teaches that the exhaust gas duct is then connected upstream of at least the last pump in the row of pumps as viewed in the main fuel flow direction from a fuel source to the combustion chamber. These exhaust gas ducts may however also be connected further upstream even upstream of the first turbo pump of a row of turbo-pumps, for example.
The bypass fuel supply line which feeds fuel to the gas generator can be connected to different fuel sources. For example it is possible that the auxiliary bypass fuel supply lines for the gas generator branch off from the main fuel supply lines for the rocket engine. This branching off of some of the fuel to the gas generator can connect to the main fuel supply line or lines at any suitable location. It is however preferred, that the bypass fuel supply lines for the gas generator are branching off from the main fuel supply line or lines downstream of the turbo pumps. This combination of features according to the invention makes sure that the gas generator is supplied with fuel which is already under a high pressure downstream of the pump or pumps whereby it is possible to achieve higher pressures in the gas generator itself and in the exhaust gases from the gas generator. Higher pressure exhaust gases facilitate on the one hand the return of the exhaust gases into the main fuel flow and on the other hand reduce the impulse losses of the rocket engine.
In another embodiment, the invention provides that a first fuel bypass line is at least partially supplied with fuel from a separate fuel source for the gas generator, for example a separate fuel tank. Using a separate fuel tank for the fuel supply to the gas generator has the advantage that the gas generator may be operated with a different fuel composition or even basically with other fuels altogether than the combustion chamber of the rocket engine. Such a rocket engine thus corresponds to a three-component fuel system or a multi-component fuel system.
In another preferred embodiment the invention provides at least one compressor stage such as a booster pump or jet pump arranged in the main fuel supply line upstream of the respective turbo pump. This feature helps increasing the pressure in the main liquid fuel flow upstream of the turbo pumps particularly in order to make possible the pressure conditions that are required for introducing the exhaust gases from the gas generator upstream of the turbo pumps. The exhaust gas duct from the gas generator is connected to or merges into the main fuel supply line in the area of the compressor stage in the main fuel supply line. The turbo pumps in the main fuel supply line or lines are preferably double flow pumps in order to simultaneously improve the cavitation characteristics. The introduction of the exhaust gases of the gas generator under excess pressure into the main fuel lines also has a beneficial influence on the cavitation characteristic.
An evaporation of the liquid fuel in the main fuel flow can be avoided by a suitable adaptation of the pressure and temperature states in the area where the exhaust gas duct or ducts lead into the main fuel flow. The cavitation characteristic can be further improved by a respective construction of the blades of the turbo pump to counteract an incomplete condensation of the exhaust gases even in spite of a rapid intermixing of the exhaust gas of the gas generator with the main fuel flow.
The condensation of the exhaust gas from the gas generator in the main fuel flow can be further facilitated by reducing the temperature of the exhaust gas prior to the introduction of the exhaust gas into the main fuel flow. For this purpose, for example, a heat exchanger is arranged for cooperation with the exhaust gas duct for cooling the exhaust gas of the gas generator. Such a heat exchanger may for example be provided in the form of a thermal contact between the exhaust gas duct and the main fuel flow line. For example, the exhaust gas duct may be provided in the form of a coil with a number of turns around the main fuel flow line whereby the respective thermal contact is provided between the exhaust duct and the main fuel supply line. The cool main fuel flow thereby reduces the temperature of the gas generator exhaust gas. Another possibility for reducing the temperature of the gas generator exhaust gases is achieved by mixing a portion of the main fuel flow with the exhaust gas with the aid of a further bypass fuel line or fuel branch line which feeds cool fuel from the main line into the exhaust gas duct or ducts. The exhaust gases of the gas generator are thus cooled before they are introduced into the main fuel flow by the mixing with a certain proportion of the cool main fuel whereby the temperature of the exhaust gases can be notably lowered and a condensation of the exhaust gas at the introduction of the exhaust gas, that follows the mixing, into the main fuel flow is facilitated.
In order to facilitate the ignition of the rocket engine during the starting phase the invention provides that the exhaust gas duct comprises duct branches with a switch over or two-way valve whereby a first exhaust duct branch is connected upstream of the turbo pumps to the main fuel flow line while a second exhaust gas branch is connected downstream of the turbo pumps to the main fuel flow line. During the starting phase the switch over valve is so controlled in closed loop fashion that the gas generator exhaust gases are first introduced only downstream of the turbo pumps into the main fuel flow and after starting the valve is switched over to introduce the exhaust gas into the main fuel flow upstream of the turbo pumps. Once the operational state of the rocket engine has stabilized itself sufficiently, the valve is switched over so that the gas generator exhaust gases are introduced into the main fuel flow upstream of the turbo pumps.
The second exhaust gas branch is switched off during normal operations of the liquid fuel rocket engine.